Convertible aircraft with tilting rotors

ABSTRACT

Improving convertible aircraft of the aforementioned type mainly in terms of the active control of vibrations, while at the same time avoiding the addition of significant additional masses for essentially countering the excitations caused by the rotors and, secondarily, of making such an aircraft easier to control in terms of roll in airplane mode. In this secondary objective, these improvements aim to allow simplifications to the maneuvering and command means for controlling the aircraft in terms of roll in airplane mode and preferably, at the same time, a simplification of the command and maneuvering means performing the functions of augmenting lift and reducing offsetting power. The invention also relates to improvements made to the convertible aircraft of the aforementioned type to give them the ability to land in airplane mode (without converting beforehand from airplane configuration to helicopter configuration). This possibility makes it possible to reduce the critical nature of the rotor tilt mechanisms in that landing remains possible without damage to the aircraft, whatever the position of the rotors. Furthermore, if both engines fail in airplane mode it is advantageous, particularly from the safety point of view, to be able to make a descent in airplane mode and in gliding flight and to land without having to tilt the rotors into helicopter mode before setting down, hence reducing the workload on the crew.

The invention relates to improvements made to convertible tilt-rotoraircraft which can travel in airplane mode and in helicopter mode, andof the type comprising at least one fuselage, at least one empennagewith at least one stabilizer, a fixed wing structure comprising at leasttwo wings extending laterally on each side of said fuselage and, inhelicopter mode, a rotary wing structure comprising at least two rotors,each of which is supported and driven in rotation by a respective one oftwo drive nacelles each supported by a respective one of the two fixedwings extending from the fuselage as far as the corresponding nacelle,each rotor being mounted so that it can tilt with at least one frontpart, which supports said rotor, of the corresponding nacelle on thecorresponding fixed wing and about an axis of tilt which is roughlytransversal with respect to the fuselage, so as to switch betweenhelicopter mode and airplane mode, in which mode the rotors act aspropellers.

As is known, these convertible aircraft can operate in helicopter modeor configuration, particularly for landings and take-offs, during whichthe rotors rotate above the fixed wings, about axes that are roughlyvertical so as to provide the aircraft with lift, and in airplane modeor configuration, in which the rotors are tilted with respect to thefixed wings so as to operate as propellers.

Each rotor has its shaft connected by a respective transmission to arespective engine, the transmission and the engine being housed in thecorresponding nacelle supported by the corresponding fixed wing, aninterconnecting shaft connecting the two transmissions, so that the tworotors can be driven in rotation by either one of the two engines,should the other engine fail.

U.S. Pat. No. 5,054,716 describes a first example of a convertibleaircraft of this type, in which each of the rotors, together with itsoperating means, the corresponding engine and the correspondingtransmission, constitutes a tilting assembly housed, with the exceptionof the rotor blades and hub, in a nacelle mounted so that the entirething can pivot, cantilever fashion at the tip of a corresponding fixedwing.

An architecture such as this has numerous disadvantages, mentioned inpatent application FR 99 03735, which describes another architecture ofa convertible aircraft of this type, overcoming the aforementioneddrawbacks and in which each transmission comprises a front reductiongear assembly, driving the rotor in rotation, and a rear reduction gearassembly engaged with the corresponding front reduction gear assemblyand connected to the corresponding engine, and to the interconnectingshaft connecting the two transmissions. Each of the two nacelles isarticulated and comprises a front part, mounted so that it can tiltabout the axis of tilting on a fixed rear nacelle part fixed to thecorresponding fixed wing and in which rear nacelle part thecorresponding engine and at least part of the rear reduction gearassembly of the corresponding transmission are housed. The frontreduction gear assembly and the corresponding rotor shaft are housed inthe tilting front nacelle part and are mounted so that they can tiltwith said front part with respect to said rear part and saidcorresponding wing.

Whatever their architecture—nacelles which tilt in full with the rotors,or articulated nacelles only the front parts of which tilt with therotors with respect to the fixed wings—the convertible aircrafts of theaforementioned type present, in terms of vibration control, new problemswhich are far removed from those presented by helicopters. Specifically,the special architecture of convertible aircraft, with rotors andpossibly engines that can tilt at the wing tips, makes it very difficultto filter vibrations by inserting anti-resonant elements as is done onhelicopters.

By contrast, the use of vibration filtration systems using resonators orforce generators driven by computers is known, but these systemsgenerally involve spring-mass assemblies which are very penalizing interms of weight.

Furthermore, in a convertible aircraft of the aforementioned type, rollcontrol in airplane mode is normally provided by the use of orientablecommand and/or control surfaces each mounted to pivot about an axissubstantially transversal to the aircraft, along the trailing edge ofeach of the fixed wings, and these orientable surfaces are also used toprovide the aircraft with additional lift at low speed in airplane modeand to reduce the offsetting power of the wing in helicopter mode.

As roll control requires swift dynamics, each fixed wing has arelatively high number of such command and/or control surfaces along itstrailing edge, and these orientable surfaces need to be able to becommanded by a great many activators and may be connected by complexmechanical links allowing, in particular, negative turns (upward) andasymmetric turns of these surfaces across the two wings, while at thesame time limiting the risk of a runaway roll situation associated withthe asymmetric turnings of these orientable surfaces which are known asflaperons because they can perform the functions of flaps and ofailerons, and therefore of lift-augmenting and warping control surfaces.U.S. Pat. No. 5,094,412 describes means of complex structure and complexcontrol of such flaperons, also known as elevons, for a convertibleaircraft according to the aforementioned U.S. Pat. No. 5,054,716.

The problem underlying the invention is that of improving convertibleaircraft of the aforementioned type mainly in terms of the activecontrol of vibrations, while at the same time avoiding the addition ofsignificant additional masses for essentially countering the excitationscaused by the rotors and, secondarily, of making such an aircraft easierto control in terms of roll in airplane mode.

In this secondary objective, these improvements aim to allowsimplifications to the maneuvering and command means for controlling theaircraft in terms of roll in airplane mode and preferably, at the sametime, a simplification of the command and maneuvering means performingthe functions of augmenting lift and reducing offsetting power.

The invention also relates to improvements made to the convertibleaircraft of the aforementioned type to give them the ability to land inairplane mode (without converting beforehand from airplane configurationto helicopter configuration). This possibility makes it possible toreduce the critical nature of the rotor tilt mechanisms in that landingremains possible, without damage to the aircraft, whatever the positionof the rotors.

Furthermore, if both engines fail in airplane mode, it is advantageous,particularly from the safety point of view, to be able to make a descentin airplane mode and in gliding flight and to land without having totilt the rotors into helicopter mode before setting down, hence reducingthe workload on the crew.

However, the ability of a convertible to land in airplane mode isusually accompanied by a reduction in the size of the rotors, andsometimes in the replacement of the rotors with airscrews, as is thecase in a third architecture of convertible aircraft known as the tiltwing aircraft, in which the aircraft wings pivot in full or in partabout the axis of tilting with the nacelles that they support.

Now, reducing the size of the rotors has known consequences on theperformance of an aircraft of the VTOL (Vertical Take-Off and Landing)type. These consequences are, in particular:

degradation in the performance in hovering flight and at low speed,because the lift effectiveness of a rotor decreases rapidly with itssize, thus eliminating the apparent gain achieved by effacing the wingsunder the rotors in embodiments with wings tilting with the nacelles androtors, for example,

increasing the external noise, which is associated with the increase inthe load at the rotor disks (maximum mass divided by the area of therotor disks), and

degradation in the ability to windmill or autorotate which is associatedwith the increase in the load at the rotor disks.

The second problem underlying the invention is that of overcoming thesedrawbacks by the use of rotors of relatively large size, sized tooptimize performance in hovering flight, with cyclic pitch control(which does not generally exist in tilt wing embodiments) and collectivepitch control for good convertible-aircraft behavior in helicopter modeand during conversion, such large-sized rotors being tiltable withrespect to fixed wings, so as to keep the advantages inherent in thistype of wing structure, particularly so as to limit the drag in forwardflight in helicopter mode (to improve performance on take-off in theevent of an engine failure) with respect to a completely or partiallytilting wing, and which allows good behavior during conversion. Inaddition, the wing structure of the convertible aircraft is configuredto allow landing in airplane mode in spite of the presence of relativelylarge-sized rotors.

With the effect of solving the first problem underlying the inventionand as explained hereinabove, the convertible aircraft according to theinvention is characterized in that each fixed wing is extended,substantially in the direction of its span and toward the outboard sideof the corresponding nacelle with respect to the fuselage, by at leastone outboard wing portion, at least part of which pivots, independentlyof the rotor and of at least the front part of the correspondingnacelle, about an axis of articulation substantially transversal to theaircraft and constitutes an orientable command and/or control surface,whose pivotings about the axis of articulation are commanded, at leastat a frequency of the order of KbΩ, where b and Ω are, respectively, thenumber of blades and the frequency of rotation of each rotor, and K is awhole number at least equal to 1, by at least one driven actuator so asto at least attenuate, at the fuselage, the fixed wings and theempennage(s) and stabilizer(s), at least the vibrations generatednaturally by the rotation of each rotor.

Aside from controlling the vibrations generated by the rotors in normaloperation, or natural vibrations inherent in the rotors and generated inrotating axis, the orientable and outboard wing tip command and/orcontrol surfaces can be used to provide active control of the vibrationsproduced by either or both of the phenomena of whirl flutter and tailshake, the former of which, it will be recalled, is an aeroelasticinstability arising from the looping between a rotor and thecorresponding wing at high speed in airplane mode, while the lattercorresponds to vibrations of the tail boom or of the rear parts of thenacelles and of the fuselage of a convertible aircraft, said tail boomor rear parts of nacelles and of the fuselage being excited by the washof the rotors at frequencies of a few hertz and often close to 4 Hz.This active control makes it possible to reduce the stiffness, andtherefore the mass, of each wing for a given maximum speed and/or toincrease the maximum speed of the aircraft.

To this end, the pivoting of the orientable command and/or controlsurface is transiently commanded by the driven actuator at a frequencybelow KbΩ and of the order of 4 to 6 Hz so as to counter the whirlflutter phenomenon, or of the order of a few Hz, generally of the orderof 4 Hz, so as to counter the tail shake phenomenon.

When integrated into an active anti-vibration system, the outboardorientable command and/or control surfaces, arranged as elevonsoperating as little driven ailerons or active aerodynamic flaps,situated at the wing tips, generate aerodynamic forces which are used tocounter the excitations generated by the rotors, avoiding the additionof additional masses.

A convertible aircraft of the type set out hereinabove is thus equippedwith a self-adaptive anti-vibration system based on the outboardorientable command and/or control surfaces or outboard elevons operatingas driven ailerons and/or as active aerodynamic flaps at the wing tips,and the turn angle, and therefore incidence of which is driven by atleast one computer commanding at least one maneuvering actuator, firstactuator being controlled at the frequency bΩ, any second actuator beingcontrolled at the frequency 2 bΩ, any third actuator being controlled atthe frequency 3 bΩ, etc., so as to generate aerodynamic forces directedagainst the excitation forces of the rotors and thus making it possibleto minimize the level of vibration in the fuselage, the empennage(s) andstabilizer(s), the fixed wings and any fixed rear parts of the aircraftnacelles that there might be, this anti-vibration system beingparticularly well suited to operation in airplane mode.

To this end, and advantageously, the actuator is an excitation ramslaved in movement, maneuvering the orientable command and/or controlsurface against the action of static and dynamic tuning elastic meansand driven automatically by at least one active and self-adaptivevibration control computer which drives the ram on the basis of signalsreceived from sensors, particularly load, accelerometer and gyrometersensors, arranged at least at predetermined points on the fuselageand/or the rotors and/or the empennage(s) and stabilizer(s), inparticular.

Advantageously, the elastic means absorb the static forces of theorientable surface and, in dynamic terms, their stiffness is coupled tothe inertia of the moving assembly comprising at least said orientablesurface and the moving parts of the ram so as to create a second-orderresonant system, the resonant frequency of the moving assembly beingtuned to the excitation frequency of the ram, which makes it possible toconsiderably reduce the control forces, and therefore the size of theram.

In practice, the moving assembly has a resonant frequency${f = {\frac{1}{2\quad \pi}\sqrt{\frac{k}{I}}}},$

where k is the stiffness of the elastic means (25) and I is the inertiaof the moving assembly and the excitation frequency of the ram isnormally tuned to bΩ such that bΩ is substantially equal to f.

When the two rotors of the convertible aircraft are three-bladed rotors,and given the nominal rotational speed of the rotors, the excitationfrequency of the ram is normally tuned substantially to a frequency ofabout 20 Hz.

In addition, in order to counter the phenomena of tail shake and whirlflutter, the excitation frequency of the ram is advantageouslytransiently tuned to a frequency of about 4 Hz to about 6 Hz when saidsensors detect signals that bear witness to at least one of these twophenomena, then, once the phenomenon has been attenuated or hasdisappeared, the excitation frequency of the ram is tuned back tosubstantially the frequency bΩ.

Such an outboard orientable surface (outboard of a nacelle) can alsoreceive a differential command, with respect to the orientable surfaceoutboard of the other nacelle, and operate as an aileron commandingwarping and allowing the aircraft to be controlled in terms of roll inairplane mode, with the required swift dynamics, the roll commandafforded by the outboard orientable surfaces thus being decoupled fromthe lift-augmenting and lift reduction functions that can be carriedout, with slower dynamics, by other inboard (between the nacelles andthe fuselage) orientable command and/or control surfaces on the trailingedges of the fixed wings.

In this alternative form, the excitation ram of each elevon or outboardorientable surface is also driveable by pilot controls (actuated by thecrew of the convertible aircraft), particularly warp controls. In thiscase, command of the excitation ram by the vibration control computer isneutralized while the excitation ram is being commanded by the pilotcontrols.

EP-0 416 590 and U.S. Pat. No. 3,666,209 disclose convertible aircraftthe wing structure of which comprises aerodynamic lift-creating andpivoting surfaces outboard (along the wing span) of drive nacelles andof a fixed inboard wing portion. However, each drive nacelle pivots witha wing part about an axis of pivoting, which means that theseconvertibles have the third of the aforementioned convertible aircraftarchitectures known as the tilt wing architecture and their pivoting andoutboard aerodynamic surfaces are intended to correct variations inattitude of the aircraft about its center of gravity and are thereforecontrol surfaces activated by flight controls situated in the cockpit.Each flight control or pilot control allows the aircraft to be movedabout one of its axes of roll, pitch and yaw.

Depending on the case, these moving surfaces can be likened to aileronsallowing warping (rotation about the axis of roll) in flight in airplanemode or a movement of yaw in vertical flight in helicopter mode.

According to the present invention, the pivoting parts of the outboardwing structure portions are, unlike in EP-0 416 590 and U.S. Pat. No.3,666,209, elevons, the functions of which have been defined hereinaboveand which are self-driven to minimize, in the structure of the aircraft,vibrations which are at least of the aforementioned three types:vibrations generated by the rotors in normal operation, and by thephenomena of tail shake and whirl flutter.

The outboard orientable surfaces according to the invention can also beused to reduce the rates of descent of the aircraft with the rotorswindmilling or autorotating (in the event of a failure of the twoengines), contributing to the lift of the aircraft if these orientableoutboard surfaces are directed into the wind.

Furthermore, the presence of such outboard orientable surfaces has theimpact of increasing the aerodynamic elongation of the wings andtherefore of reducing the induced drag, thus improving performance inairplane mode in a climb, in cruising flight and in fineness, hencegiving a lower rate of descent in unpowered flight (engine failure).

By analogy with the embodiment of the conventional ailerons and flaps,the pivoting part or orientable command and/or control surface of eachoutboard wing structure portion may be a pivoting trailing edge elevonof a fixed and outboard wing portion, which thus constitutes the tip ofthe corresponding fixed wing, beyond the corresponding nacelle. However,in a second embodiment, each outboard wing structure portion may be anoutboard wing part that is entirely pivoting about the axis ofarticulation so that the entirety of the wing structure part outboard ofa nacelle can be arranged as an elevon pivoting about its axis ofarticulation with respect to the adjacent nacelle and with respect tothe corresponding fixed wing.

With a view to solving the second problem underlying the invention andas is set out hereinabove, the convertible aircraft according to theinvention is such that its fixed wings are high wings secured to theupper part of the fuselage, to keep the nacelles and therefore therotors at a sufficient height, guaranteeing a minimum ground clearanceof the rotors to allow landing in airplane mode, this ground clearancebeing increased and/or the diameter of the rotors being increased if thehigh wings are raised with respect to the upper part of the fuselage.

However, advantageously in addition, the fixed high wings have an upwarddihedral angle (positive dihedral angle) between the fuselage and thenacelles, which, at the same time, makes it possible to increase theground clearance and/or the diameter of the rotors still further andmakes it possible to limit the drag penalty due to the raised positionof the fixed wings above the fuselage.

Fixed wings which are raised with respect to the fuselage and with anupward dihedral angle undeniably improve the landing capability with therotors in airplane mode.

Outboard of the nacelles, the outboard wing structure portionscomprising the outboard orientable command and/or control surfaces, orarranged as such outboard orientable surfaces, may also have a positive(upward) or zero (substantially horizontal) dihedral angle, butadvantageously, in order to at least partially compensate for anydisadvantageous aerodynamic effects of the positive dihedral angle ofthe fixed wings, the outboard wing structure portions may have anegative (downward) dihedral angle so that the fixed wing structure ofthe aircraft is substantially in the form of gull wings.

It should be noted that the characteristics relating to the raised fixedhigh wings with an upward dihedral angle and which are possiblyextended, outboard of the nacelles, by outboard wing structure portionswith a positive, zero or negative dihedral angle, can be used on aconvertible aircraft of the type set out hereinabove independently ofthe other characteristics set out hereinabove and relating to thestructure, the arrangement, the maneuvering and the control of theoutboard orientable command and/or control surfaces of the convertibleaircraft, and vice versa. Specifically, such outboard orientablesurfaces may equip the wing tips of a fixed wing structure of aconvertible aircraft, the fixed wings of which are not high wings norare they raised nor do they have a positive dihedral angle.

The invention will be better understood, and other features andadvantages of the invention will become apparent from the descriptiongiven hereinbelow with no implied limitation of some exemplaryembodiments which are described with reference to the appended drawingsin which:

FIGS. 1, 2 and 3 are schematic views respectively in side elevation,front elevation and plan view, of a first embodiment of a convertibleaircraft equipped with a gull wing fixed wing structure and withoutboard wing structure portions arranged as outboard wing parts whichpivot entirely about axes of articulation and therefore constituteelevons operating as driven ailerons and/or as active aerodynamic flaps,

FIG. 4 is a schematic partial plan view of a second exemplaryembodiment, in airplane mode,

FIG. 5 is a schematic sectional view of the outboard orientable surfaceof the example of FIG. 4 and of its maneuvering and control means,

FIG. 6 is an alternative form of embodiment of FIG. 4, and

FIG. 7 is a curve depicting, as a function of frequency, theamplification of the amplitude of the movements of the elevon withrespect to the amplitude of the excitations of the control ram, in amechanical system according to FIG. 5.

The convertible aircraft 1 of FIGS. 1 to 3 comprises a fuselage 2, ofthe airplane fuselage type, with a cockpit 3 at the front and a T-shapedempennage 4 with stabilizer(s) at the rear, and two fixed high wings 5,in this example with zero backsweep and constant chord (rectangular inplan view), extending, in the conventional way, laterally on each sideof the fuselage 2, each fixed wing 5 supporting, at its outboard endalong its wing span on the opposite side to the fuselage 2, anarticulated drive nacelle 6 fixed by its rear part 7 to thecorresponding wing 5. As an alternative, the empennage 4 may have someother geometric configuration, for example in the form of a cross or ofa V, or of twin stabilizers, or some other configuration. Theconvertible aircraft may be equipped with one or more empennages 4equipped with one or more stabilizer(s).

Each nacelle 6 comprises a streamlined front part 8, substantially inthe shape of an ogive, and mounted so that it can tilt, with respect tothe fixed rear part 7 of the nacelle 6, and therefore with respect tothe corresponding wing 5, about an axis of tilting X—X which istransversal to the fuselage 2 and more particularly perpendicular to thelongitudinal plane of symmetry passing through the longitudinal axis A—Aof the aircraft 1.

When the aircraft is in this aerodynamic configuration, the axis oftilting X—X is parallel to the plane perpendicular to the longitudinalaxis A—A of the aircraft 1 and containing the straight lines that passthrough the centers of aerodynamic pressure of the wings 5, the axis oftilting X—X being slightly behind the lines of the centers of pressuresituated a distance from the leading edge of the wings 5 which is about25% of the value of the chord of the wings 5, toward their trailingedge.

On the tilting front part 8 of each nacelle 6, a rotor, for example athree-bladed or four-bladed rotor, depicted diagrammatically in the formof the rotor disk 9, is mounted so that it rotates about its axis andalso tilts about the axis X—X with the corresponding nacelle front part8, each rotor 9 having a shaft connected, for its rotational drive, by atransmission to a turbine engine unit fixed in the fixed rear part 7 ofthe corresponding nacelle 6, in an arrangement described morespecifically in FR 99 03735, to which reference can be made for furtherspecifics on that subject, and which is incorporated into this text byreference.

To drive the two rotors 9 when one or other of the turbine engine unitsis defective, and regardless of the inclination of the rotors 9 and ofthe tilting front parts 8 of the nacelles 6 about the axis of tiltingX—X, the two transmissions are joined together by an interconnectingshaft depicted diagrammatically as 10 in FIG. 3. Reference may also bemade to FR 99 03735 for further details on the various possibleconnections of the interconnecting shaft 10 to the two transmissions ofthe rotors 9. The shaft 10 essentially consists of two straight parts,each extending between the fuselage 2 and a respective one of thenacelles 6, along the entire span of the corresponding fixed wing 5,this shaft part 10 being substantially parallel to the planeperpendicular to the longitudinal axis A—A and passing through the axisof tilting X—X, but offset with respect to that plane, for exampletoward the rear of the wings 5 (see FIG. 3), the two straight parts ofthe interconnecting shaft 10 being coupled together on the top of thefuselage 2, and coupled to an accessories drive unit housed in a raisingbox structure 11 via which the fixed high wings 5 are secured to theupper part of the fuselage 2, being raised with respect to this fuselage2.

In order for each rotor 9, tilting with the tilting front part 8 of thecorresponding nacelle 6, to be able to be driven in rotation about itsaxis by the turbine engine unit housed in the fixed rear part 7 of thisnacelle 6 or by the interconnecting shaft 10 driven off the otherturbine engine unit housed in the nacelle 6 of the other wing 5, eachtransmission, which has a free wheel to neutralize the differences inrotational speed of the engines, comprises a transmission front part orfront reduction gear assembly, which drives the shaft of the rotor 9 andis mounted to tilt with it about the axis of tilting X—X, inside thetilting front part 8 of the nacelle 6, the front reduction gear assemblyremaining in constant mesh with a rear transmission part or rearreduction gear assembly which does not tilt, in constant mesh with apower take-off at the corresponding end of the interconnecting shaft 10and with an output shaft of the corresponding turbine engine unit. Therear reduction gear assembly which does not tilt, occupies a fixedposition with respect to the turbine engine unit, to the rear part 7 ofthe corresponding nacelle 6, to the interconnecting shaft 10 andtherefore to the wings 5, and is partially housed in the fixed rear part7 of the nacelle 6 and possibly in at least one of the streamlinedconnectors or fairings between the fixed rear nacelle part 7 and thecorresponding wing 5.

The convertible aircraft 1 is thus equipped with two tilt rotors each ofwhich can be tilted between a helicopter configuration in which therotors are labeled 9′ on the front nacelle parts which are labeled 8′,and in which each rotor 9′ is driven about a substantially vertical axisof rotation for operation of the aircraft in helicopter mode, and anairplane configuration in which each rotor labeled as 9 (in FIGS. 1 to3) at the front of a nacelle front part labeled as 8 behaves as apropeller driven in rotation about an axis substantially aligned withthe direction of flight, for operation of the aircraft in airplane mode.

In practice, the tilting front part 8 of each nacelle 6 tilts via acentral rear portion between two front lateral extensions of the fixedrear part 7 of this nacelle 6, as described in FR 99 03956.

To ensure good supply of air to the engine, fixed in this fixed rearpart 7, the latter has a lower air intake 12 which is fixed and whichopens forward under the tilting front part 8 of the nacelle 6.

The fixed wings 5, raised at 11 on the fuselage 2, also have a positivedihedral angle (see FIG. 2), that is to say that each of them is raisedupward, with respect to their raised connection 11 to the fuselage 2, onthe side of their end supporting the nacelle 6. The nacelles 6, andtherefore the rotors 9, are thus raised up above the ground so thatsufficient ground clearance remains between the rotors 9 and the ground,even when the rotors 9 are of relatively large diameter, to allow theaircraft 1 to land with the rotors 9 in airplane configuration,particularly in the event of failure of both engines, the rotors 9 thenwindmilling or autorotating, declutched from the engines, but theaircraft can also be landed in airplane mode with rotors 9 driven by theengines or stationary.

Adopting raised wings 5 with a positive dihedral angle therefore allowsthe aircraft 1 to be landed without damage regardless of the position ofthe rotors 9, and this possibility reduces the criticality of thesystems used for tilting the rotors 9. In the event of failure of thetwo engines in airplane mode, the aircraft can descend in gliding flightand land without having to tilt the rotors 9 into helicopter mode priorto landing.

Note that the space freed up in the raising box structure 11 above thefuselage 2 for raising the wings 5 allows additional accessories such asalternators, hydraulic pumps, etc., driven by the intermediate gearboxhoused in this raising box 11 to be housed and provides theinterconnection of the two parts of the interconnecting shaft 10. Thismakes incorporating the engine and the rear transmission part into eachnacelle 6 simpler and makes it possible to reduce the size of thenacelles 6.

The positive dihedral angle of the wings 5 not only makes it possible toraise the rotors 9 high enough to allow landing in airplane mode, butalso makes it possible to use rotors 9 of sufficiently large size thatthey can be sized to optimize performance in hovering flight, withcyclic and collective pitch control for good behavior in helicopter modeand when converting between the airplane and helicopter modes.

Furthermore, the penalty in terms of drag due to the presence of theraising box 11 allowing the raised position of the wings 5 above thefuselage 2 is limited by the use of wings 5 that have a positivedihedral angle.

These advantages stemming from wings 5 with a positive dihedral angleadd to those afforded by the fixed wings, namely the limitation of dragin forward flight in helicopter mode by comparison with a partially orfully tilting wing, and the obtaining of good behavior duringconversion, because the wings 5 do not stall. This limitation in drag inforward flight in helicopter mode makes it possible to improveperformance on take-off in the event of an engine failure.

According to another feature specific to the convertible aircraft 1 ofthe invention, the aircraft 1 wing structure comprises, substantially inthe continuation of each wing 5, along its span and outboard of thecorresponding nacelle 6, that is to say on the opposite to the fuselage2, an outboard wing structure portion 13, at least part of which pivotsabout an axis of articulation Y—Y substantially transversal to theaircraft 1, and preferably contained in a plane perpendicular to thelongitudinal axis A—A of the aircraft 1, that is to say in a planeparallel to the axis of tilting X—X of the rotors 9 with at least thefront parts 8 of the nacelles 6 and the front parts of thetransmissions.

In the example of FIGS. 1 to 3, each outboard wing structure portion 13is an outboard wing part that pivots in its entirety about the axis ofarticulation Y—Y. Each of the two outboard wing parts 13, ofsubstantially trapezoidal shape in plane view (see FIG. 3), and thestraight leading and trailing edges of which converge toward one anotherlaterally toward the outboard end and are connected by a small tipfairing 14, constitutes a command and/or control surface which isorientable, and the pivotings of each about its axis of articulation Y—Yare controlled by an actuator, such as a ram, driven by a computer whichmay be incorporated into the pilot control computers. These orientablesurfaces 13 perform the functions of driven ailerons and/or of activeaerodynamic flaps, situated at the tips of the wings 5, beyond thenacelles 6, and therefore constitute evelons which can operate as pitchcontrol surfaces and as warp control surfaces. The axis of articulationY—Y makes it possible to control the incidence of each elevon 13, andits maneuvering ram (not depicted) is slaved in movement and drivenconstantly by the computer to control the turning of the elevon 13.

Operated as ailerons, the orientable elevons 13, situated outboard ofthe nacelles 6, allow the aircraft 1 to be controlled in terms of rollin airplane mode, this roll control, which demands swift dynamicresponse, thus being decoupled from the lift-augmenting andlift-reducing functions which are performed by orientable command and/orcontrol surfaces mounted along the trailing edges of the fixed wings 5between the nacelles 6 and the fuselage 2 and which constitute inboardailerons and/or flaps 15, in low number, pivoting about axes ofarticulation Y′—Y′ also substantially transversal with respect to thefuselage 2, for example two ailerons and/or flaps 15 per wing 5 (seeFIGS. 2 and 3).

As a result, this lower number of inboard orientable surfaces 15 can becontrolled by a lower number of slow dynamic actuators, which aretherefore simple and economical. In addition, the structure of theinboard orientable surfaces 15 can be simplified, and their number mayeven be reduced to one surface 15 per wing 5, the need for negativeturning (orientation upward) disappearing as does the need forasymmetric turning between the two wings 5. In this alternative form andunder these conditions, a mechanical link between the two orientablesurfaces 15 makes it possible in a simple way to cover the risk of theaircraft 1 running away in terms of roll in response to asymmetricturning of the orientable surfaces 15, operating simply as flaps, butalso being able, more generally, to operate as elevons.

In addition to the control of the roll of the aircraft 1 in airplanemode, which is secondary, the outboard orientable surfaces 13 (oroutboard elevons 13) are used primarily as automatically driven activeaerodynamic flaps, to provide active vibration control, particularlycontrol of natural vibrations inherent to the rotors 9 and generated inrotating axis, and vibrations associated with the phenomena known astail shake and whirl flutter, the latter being essentially anaeroelastic instability arising from the coupling between the rotors 9and the wings 5 at high speed in airplane mode. This active control,described more specifically hereinbelow with reference to FIGS. 4 and 5,consists in developing, as depicted schematically in the left-hand partof FIG. 2, aerodynamic forces F2 on the elevons 13 to counter theexcitations F1 brought about by the rotors 9 in particular, themaneuvering ram of each elevon 13 constantly and automatically drivingthe fluctuations of incidence of the latter, which generates aerodynamicforces by receiving orders from the computer slaving the ram in movementand calculating the command orders from, for example, signals fromaccelerometer and/or gyrometer sensors and for load sensors situated atvarious points on the convertible such as determined points on thefuselage 2 and/or on the rotors 9 and/or on the empennage(s) andstabilizer(s) 4. Each outboard elevon 13 may be driven in incidenceabout its axis of articulation Y—Y against the action of elastic meanswhich return the elevon 13 to a neutral position in the absence ofloading on the part of the maneuvering ram, these elastic meansfurthermore having two functions, which are those of absorbing thestatic forces of the elevon 13, so as to relieve the correspondingmaneuvering ram, and so that dynamically the stiffness of the elasticmeans, coupled with the inertia of the moving parts (essentially theelevon 13 and the moving parts of its maneuvering ram) creates aresonant second-order system the resonant frequency of the moving partsbeing tuned to the excitation frequency making it possible toconsiderably reduce the control forces and therefore the size of themaneuvering ram.

This then produces an active self-driven and self-adapting antivibration system based on the injection of additional aerodynamic forcesintroduced by the control of the outboard elevons 13 to minimize thevibrations in all the non-pivoting parts of the aircraft 1 especiallythe fuselage 2, the empennage(s) and stabilizer(s) 4, the fixed wings 5and the non-pivoting rear parts 7 of the nacelles 6.

The use of the outboard elevons 13 for active vibration controltherefore makes it possible to reduce the stiffness, and hence the mass,of the wings 5 for a given maximum speed and/or to increase this maximumspeed of the aircraft 1.

The outboard elevons 13 also make it possible, when driven as flaps, andin the event of a failure of both engines, to reduce the rate of descentof the aircraft, with the two rotors 9 wind milling or autorotating, bycontributing to the lift of the aircraft 1 in airplane mode, if theelevons 13 are directed into the wind.

Finally, through their presence practically at the tips of the wings 5,the elevons 13 increase the aerodynamic elongation of the wingstructure, and this reduces the induced drag and therefore improvesperformance in airplane mode in a climb, in cruising flight and infineness (so that the rate of descent in unpowered flight is reduced).

As depicted in FIG. 2, the outboard elevons 13 have a slightly negative(downward) dihedral angle which is accentuated at their tip fairing 14,therefore in the opposite direction to the dihedral angle of the wings5, to compensate for any effects that this positive dihedral angle ofthe wings 5 might have.

It will also be understood that the use of partially tilting nacelles 6(with an engine and rear transmission part housed in the fixed rear part7 of the nacelle and the rotor 9 tilting with the front part of thetransmission and the front part 8 of the nacelle 6), allow simplestructural integration of the outboard elevons 13 to the nacelles 6.

As an alternative, the elevons or outboard wing structure portions 13may have a zero dihedral angle (may be substantially horizontal) or mayalso have a positive dihedral angle, substantially in the continuationof that of the wings 5, but these outboard wing structure surfaces 13are preferably given a slight negative dihedral angle (see FIG. 2)giving the main wing structure the appearance of gull wings.

In the schematic alternative form depicted partially in FIG. 4, it ispossible again to see each fixed wing 5′ fixed in a high and raisedposition with a positive dihedral angle on the upper part of thefuselage 2 and, in this example, with the shape in plan view of a righttrapezium with a leading edge 16 with zero backsweep and a trailing edge17 with negative backsweep and equipped with two inboard elevons 15, andwhich at its outboard end supports a nacelle 6, at least the front part8 of which tilts with a rotor 9. An outboard elevon 13′, which, in planview, has the shape of a right trapezium, the leading and trailing edgesof which substantially continue the respective leading and trailingedges 16 and 17 of the wing 5′ is mounted so that it can pivot in itsentirety, outboard of the nacelle 6, about its axis of articulation Y—Y.In this alternative form, the axis of articulation Y′—Y′ of the inboardelevons 15 is inclined with respect to the plane transversal to theaircraft passing through the axis of articulation Y—Y of the outboardelevon 13′, whereas in the example of FIGS. 1 to 3, the axes ofarticulation Y′—Y′ of the inboard elevons 15 and those Y—Y of theoutboard elevons 13 are in parallel transverse planes.

FIG. 5 schematically depicts the control of the orientation of eachoutboard elevon 13′ of FIG. 4 about its axis of articulation Y—Y. Toproduce a self-adaptive antivibration system based on the outboardelevons 13′, the turn angle of each outboard elevon 13′ is driven by acomputer 18 receiving at 19 signals from accelerometers, gyrometers andload sensors arranged at given points particularly on the fuselage 2,the rotors 9 and the empennage(s) and stabilizer(s) 4 of the aircraft.The computer 18 drives an excitation ram 20, which is a linear ram,slaved in movement, the cylinder 21 of which bears against a fixed point22 of the fixed rear part 7, for example, of the nearby nacelle 6, or ona fixed point 22 of the wing 5′, while the piston and the rod 23 of theram 20 drive a small lever 24 secured to the elevon 13′ in rotationabout the axis of articulation Y—Y and against the action of a staticand dynamic tuning spring 25 also urging the lever 24 via one end andresting via its other end on a fixed point 26 of the structure of thenearby nacelle 6 or of the wing 5′. The back and forth linear movementsof the ram rod 23, shown schematically by a double-headed arrow underthe ram 20, are thus converted into back and forth rotations of theelevon 13′ about its axis of articulation Y—Y in the direction of thecurved double-headed arrow at the front of this elevon 13′ in FIG. 5.Thus, the turn angle of the outboard elevon 13′ driven by the computer18 and the ram 20, makes it possible to generate aerodynamic forces(such as F2 in FIG. 2) directed against the excitation forces of therotors 9 (such as F1 in FIG. 2). It is thus possible to minimize thelevel of vibration in the fuselage, the empennage(s) and stabilizer(s)and the fixed wings of the aircraft, particularly in airplane mode.Schematically, the excitations of the rotors 9 are countered by the lifton the outboard elevons 13′ or 13 in the examples of FIGS. 1 to 4. Asmentioned above, the spring 25 absorbs the static forces of the elevon13′ so as to relieve the ram 20 and, dynamically, the stiffness of thespring 25 coupled with the inertia of the moving assembly, comprisingmainly the elevon 13′ with its lever 24, the piston and the rod 23 ofthe ram 20 and the spring 25, creates a resonant second-order system theresonant frequency of this moving assembly being tuned to the excitationfrequency of the ram 20, this making it possible to reduce the controlforces delivered by the ram 20 and therefore the power and bulk thereof.

It is appropriate to note that the control of the ram 20 by thevibration control computer 18 is neutralized when the ram 20 is beingcontrolled by the pilot controls, so as to drive the elevon 13 or 13′ asa pitch control surface or warp control surface.

Thus, by way of example, an outboard elevon such as 13 or 13′, but onewhich has a substantially rectangular shape in plan view, with a chord0.56 m long and a span of 0.25 m, and which is driven with a turn angleamplitude of ±5° is enough to act against an excitation force of 1000 N,at the maximum speed of the aircraft in airplane mode of approximately150 m/s, the estimated mass per elevon being about 2 kg.

To produce a good influence on the change in the level of vibration dueto the dynamic forces, the outboard elevons 13 or 13′ are produced insuch a way that their center of gravity is forward of their center ofrotation on the axis of articulation Y—Y, the center of rotation itselfbeing positioned at their aerodynamic center of pressure, so as to avoidmoments which are due to the aerodynamic force. This makes it possibleto minimize the driving forces.

Each outboard wing structure portion, outboard of the nacelles 6, may benot completely an orientable surface but may, on the other hand, asdepicted schematically in FIG. 6 which depicts an alternative form ofFIG. 4, be an outboard wing structure portion 27 of which a front part,along its leading edge, is a fixed outboard wing portion 28, to the rearof which an orientable part 29 is mounted so that it can pivot about theaxis of articulation Y—Y and constitute the orientable command and/orcontrol surface operating as an aileron and/or as a flap, and thereforesimilar to the elevon 13 or 13′ of the previous examples.

In the examples of FIGS. 1 to 3, 4 and 6, the outboard elevons 13, 13′,29 introduce into the fixed structure of the nacelles 6 or of the fixedwings 5 or 5′, shear forces and bending moments, but little or notorsional moments, when their axis of articulation Y—Y passessubstantially through the center of twist of the fixed wings 5 or 5′.

By offsetting the axis of articulation Y—Y of the elevons 13, 13′, 29forward or backward with respect to the center of twist of the fixedwings 5 or 5′, the elevons can additionally introduce into the structureof the wings 5 or 5′ torsional moments which may be necessary if thephenomenon of whirl flutter manifests itself, and in order to counterthis phenomenon.

In all the exemplary embodiments, the elevons 13, 13′ and 29 aretherefore self-driven to attenuate, in the structure of the convertible,vibrations which are at least of the aforementioned types, mainlyvibrations resulting from the phenomena of whirl flutter and tail shakeand the vibrations generated by the rotors in normal operation.

As regards the whirl flutter phenomenon, it is understood that, if oneof the rotors 9 of the convertible is moved away from its plane ofrotation, for example under the effect of a gust of wind, the breakageof a wing element altering the stiffnesses of the wing, etc., parasiticvariations in the angle of incidence of the blades of this rotor 9 occurand this introduces additional aerodynamic forces which excite thisrotor 9 and sustain the movement. The corresponding fixed wing 5 or 5′is therefore deformed, and this may once again accentuate the movementof the rotor 9 with respect to its plane of rotation, and so on, so thatif the stiffness of the wing 5 or 5′ is not enough to dampen thesemovements and return the assembly to a position of equilibrium, thephenomenon diverges until the elements involved on the wing 5 or 5′and/or on the corresponding rotor 9 break.

In consequence, the movement of the wing 5 or 5′ is an overall bendingand torsional movement which results in particular in a predominantvertical movement of the wing 5 or 5′. This movement, and the vibrationsit generates, are precisely the movement and the vibrations that theelevon 13, 13′ or 27 is supposed to attenuate and then cancel out.

This phenomenon of instability corresponds to a frequency of the orderof 4 to 6 Hz approximately.

As regards the vibrational phenomenon known as tail shake, this is avibration of the rear parts 7 of the nacelles 6 and of the rear part ofthe fuselage 2 of a convertible aircraft, in which these rear parts areexcited by the wash from the rotors 9, and these vibrations develop atfrequencies of a few hertz and often close to 4 Hz.

As regards the vibrations generated by the rotors 9 in normal operation,or natural vibrations inherent to the rotors 9, which are generated inrotating axis, it is known that these vibrations are at three levels,namely KbΩ and (Kb±1)Ω where b and Ω are, respectively, the number ofblades and the rotational frequency of each rotor 9 and K is a wholenumber at least equal to 1. However, regardless of the nature of theexcitation of the rotor (flapping or drag) and regardless of the leveland position (with K=1, 2, 3, 4, . . . ) of the excitation frequency,the excitations in fixed plane of reference, in flapping up and down,bending and torsion, occur at frequencies KbΩ.

In other words, in fixed axis, that is to say at the fuselage 2 and atthe empennage(s) and stabilizer(s) 4 and at the fixed wings 5 and 5′ andthe fixed rear parts 7 of the nacelles 6 of the convertible, onlyvibrations in KbΩ are felt.

As the frequencies most disagreeable to man (and therefore to thepassengers and crew) are the lower ones, the priority is given to atleast attenuating, and if possible canceling vibrations at the frequencybΩ for K=1, particularly at the frequency 3 Ω when each of the tworotors 9 of the convertible is a three-bladed rotor. Given the nominalrotational speed of each rotor 9, which is the order of 400 rpm, Ω is ofthe order of 6 to 7 Hz, which means that excitation forces that have tobe canceled out are at frequencies of the order of 18 to 21 Hz, namelyabout 20 Hz, which constitutes a relatively high driving frequency.

As the resonant frequency of the moving assembly of FIG. 5, essentiallycomprising the elevon 13′ and its lever 24, the piston of the rod 23 ofthe ram 20 and the spring 25, is a frequency f such that f=1/2π{squareroot over (k/I)}

where k is the stiffness of the spring (25) and I is the inertia of thismoving assembly, the latter is produced in such a way that its resonantfrequency f is tuned by construction to the main frequency that is to bedamped, bΩ, namely about 20 Hz in the case of a three-bladed rotors 9.

First, the computer 18 operates the ram 20 in such a way that itsexcitation frequency is normally tuned to bΩ which is equal to f. Thismakes it possible to obtain the movement of the elevon 13, 13′ or 29with a minimum of force to be supplied by the ram 20. What happens is amechanical system which oscillates at its resonant frequency requires avery small addition of energy in order to set it into motion. It istherefore possible to use a small-sized ram 20, which meets the desiredgoal of saving mass and volume in particular.

The elevon 13, 13′ or 29 therefore has limited dimensions so as to havea low inertia, in order for it to be easy to move it at frequencies ofthe order of 20 Hz, which entails a system with good dynamics.

By way of example, an elevon which in plan view is rectangular, for aconvertible aircraft each of the two rotors 9 of which has a diameter ofabout 9 to 10 m, with a distance of 12 to 15 m between the axes of thetwo rotors 9, is an elevon made of carbon fiber with a mass of 4 kg andan area in plan view of 0.25 m², for example, a chord length of 0.5 mand a span of 0.5 m, this elevon being set at a mean angle of incidenceof 5°.

It is understood that this function of mainly damping the vibrations inbΩ can not in any way be performed by the inboard flaps (between thenacelles 6 and the fuselage 2) or by the tilting of an inboard wing partbecause, in both cases, the rotational inertia of these inboard elementsis very high, because of their large size, dictated in particular by thechord of the wing and the thickness of the wing (in proportion to thechord) at this point, which means that it is not possible to achievesufficiently high driving frequencies. In particular, the path band ofthe inboard flaps is too low, because it is merely of the order of 2 to3 Hz.

Returning to the self-driven and self-adapting active anti-vibrationsystem of FIG. 5, this system has the advantage that a dynamicamplification is put to good use so that the driving of the elevon 13′can be achieved with very small control forces on the ram 20, the movingassembly comprising the elevon 13′ and its lever 24, the piston and therod 23 of the ram 20 and the spring 25 already in itself, when operatingas a passive mechanical system, providing a great deal of amplificationdepicted schematically by the bell shaped curve in FIG. 7, whichrepresents the change, as a function of frequency f, of the ratio e/d0of the amplitude of the movements of the elevon 13′ to the amplitude ofthe excitations at the ram 20. This amplification ratio e/d0 is at amaximum for the resonant frequency of the moving assembly (13′-24-23-25)of FIG. 5, the construction of which is such that this natural frequencyis tuned to the frequency bΩ. The maximum effectiveness is thus obtainedfor this frequency bΩ. This amplification ratio, which is of the orderof 4 for example at the top of the curve in FIG. 7, is lower, butnonetheless always greater than 1, for the frequency fts, of the orderof 4 Hz, at which the tail shake phenomenon occurs, and for thefrequency fWF, lying substantially between 4 and 6 Hz and, for example,of the order of 5 Hz, at which the whirl flutter phenomenon occurs. Itwill be understood that the system of FIG. 5 can, with maximumeffectiveness, counter just one excitation frequency at a time, in thisinstance the frequency bΩ. However, although this system is designed toattenuate with the greatest effectiveness vibrations at the frequencybΩ, excitations which occur at other frequencies such as fts and fWF arealso attenuated, but not as effectively.

However, to attenuate and possibly eliminate the vibrations that resultfrom phenomena of whirl flutter and tail shake when the sensors onboardthe convertible detect the onset of these phenomena, the computer 18transiently commands the excitation frequency of the ram 20 so that thisfrequency is no longer tuned to the normal operating frequency at bΩ,but to a frequency from about 4 Hz to about 6 Hz, in the case of whirlflutter, or to a frequency of the order of 4 Hz in the case of tailshake, the computer 18 then setting the excitation frequency of the ram20 back to the frequency bΩ as soon as the whirl flutter or tail shakephenomenon has been sufficiently attenuated or even eliminated,something that the computer 18 can determine from information receivedat 19 from the in-board sensors.

Likewise, when the rotational frequency Ω of the rotors 9 varies,something which is also detected by the on-board load, accelerometricand gyrometric sensors, the computer 18 can adjust the excitationfrequency of the ram 20 to tune it once again to the new frequency bΩthus obtained. The computer 18 thus allows adaptation to the variations,generally limited in amplitude, of the rotational frequencies of therotors 9.

The elevons 13, 13′, 29 are thus commanded at a frequency that issubstantially tuned to the frequencies of the vibrational phenomena thatare to be attenuated, or even eliminated, by the computer 18 receivingsignals 19, identifying the vibrational regimes, from sensors mounted atvarious points on the convertible and sensitive to the excitation forcesapplied in particular to the fuselage 2 from the two rotor 9—elevon 13,13′ or 29 assemblies which are subjected to various aforementionedaeroelastic and vibrational excitations. In general, the load,accelerometric and gyrometric sensors informing the computer 18 may bearranged at any point on the convertible aircraft. The two elevons 13,13′ or 29 are driven together by the same computer 18 until aconfiguration is obtained which minimizes the measured level ofvibration.

Arranging the elevons 13, 13′ and 29 at the wing tips, that is to say inthe direction of the span outboard of the nacelles 6 carried at the endsof the fixed wings 5 or 5′ seems to be optimum for the followingreasons:

the essential excitations come from the rotors 9, these themselves beingat the ends of the fixed wings 5 or 5′ which means that the closer theresultants of the lift of the elevons 13, 13′ or 29 are to theexcitation forces (so that there is therefore no resultant moment on thewings 5 or 5′), the easier these essential excitations are to counter,

arranging the elevons 13, 13′ or 29 outboard of the nacelles 6 makes itpossible to avoid unfavorable interactions on the wings 5 or 5′themselves or with the control flaps 15 of these wings, whereas suchinteractions, particularly turbulent interference, would be induced onthe flaps 15 by an elevon driven at high frequency and arranged inboardof the nacelles 6 (between the nacelles 6 and the fuselage 2),

it is at the elevon 13, 13′ or 29, outboard of a fixed wing 5 or 5′-nacelle 6 assembly, that the modal deformation of this assembly is thegreatest; at constant load, it is therefore at this point that theeffectiveness of a system involving an elevon such as 13′ and a springsuch as 25 in FIG. 5 is at its maximum; in other words, as the work ofthe external forces is equal to the product of the load and thedisplacement (or rotation) for angular variation, at constant load, themaximum effectiveness is obtained as a result of the deflections and/ordynamic rotations at the outboard end of the wing-nacelle assemblies.

It is thus possible to equip the convertible aircraft with active andself-adapting antivibration systems, offering optimum adaptationcapability in flight in airplane mode, regardless of the mass,centering, structural dispersion and rotational speed conditions. Thisis because, from the signals from the acceleration, gyrometer and/orload sensors in the fuselage 2 and on the rotors 9, one or morecomputers such as 18 formulates (formulate) a command driving rams suchas 20 which maneuver the outboard elevons such as 13, 13′ or 29, whichgenerate aerodynamic forces intended to counter the vibrations. Theforces delivered by the elevons, maneuvered by the rams, are constantlyadjusted according to the level of vibration, so that this level isminimized within the meaning of a given criterion, for example aleast-squares or some other criterion.

Of course, the invention set out hereinabove is not restricted toconvertible aircraft with articulated drive nacelles as described in FR99 03735 and FR 99 03956, but also applies to convertible aircraft withdrive nacelles which tilt in full with the rotors, as described in U.S.Pat. No. 5,054,716.

What is claimed is:
 1. A method for the active control of the vibrationsof a convertible tilt-rotor aircraft which can travel in airplane modeand in helicopter mode, and comprising at least one fuselage, at leastone empennage with at least one stabilizer, a fixed wing structurecomprising at least two wings extending laterally on each side of saidfuselage and, in helicopter mode, a rotary wing structure comprising atleast two rotors, each of which is supported and driven in rotation by arespective one of two drive nacelles each supported by a respective oneof the two fixed wings extending from the fuselage as far as thecorresponding nacelle, each rotor being mounted so that it can tilt withat least one front part, which supports said rotor, of the correspondingnacelle on the corresponding fixed wing and about an axis of tilt whichis substantially transversal with respect to the fuselage, so as toswitch between helicopter mode and airplane mode, in which mode therotors act as propellers, each fixed wing being extended, substantiallyin the direction of its span and toward the outboard side of thecorresponding nacelle with respect to the fuselage, by at least oneoutboard wing portion, at least part, of which pivots, independently ofthe rotor and of at least the front part of the corresponding nacelle,about an axis of articulation substantially transversal to the aircraftand constitutes an orientable surface which is at least one of a commandsurface and a control surface, whose pivotings about the axis ofarticulation are commanded, by at least one driven actuator, the methodcomprising the step of commanding the pivotings of said orientablesurface at least at a frequency of the order of KbΩ, where b and Ω are,respectively, the number of blades and the frequency of rotation of eachrotor, and K is a whole number at least equal to 1, so as to at leastattenuate, at the fuselage, the fixed wings and the empennage andstabilizer, at least the vibrations generated naturally by the rotationof each rotor.
 2. A convertible tilt-rotor aircraft which can travel inairplane mode and in helicopter mode, and comprising at least onefuselage, at least one empennage with at least one stabilizer, a fixedwing structure comprising at least two wings extending laterally on eachside of said fuselage and, in helicopter mode, a rotary wing structurecomprising at least two rotors, each of which is supported and driven inrotation by a respective one of two drive nacelles each supported by arespective one of the two fixed wings extending from the fuselage as faras the corresponding nacelle, each rotor being mounted so that it cantilt with at least one front part, which supports said rotor, of thecorresponding nacelle on the corresponding fixed wing and about an axisof tilt which is substantially transversal with respect to the fuselage,so as to switch between helicopter mode and airplane mode, in which modethe rotors act as propellers, each fixed wing being extended,substantially in the direction of its span and toward the outboard sideof the corresponding nacelle with respect to the fuselage, by at leastone outboard wing portion, at least part of which pivots, independentlyof the rotor and of at least the front part of the correspondingnacelle, about an axis of articulation substantially transversal to theaircraft and constitutes an orientable surface which is at least one ofa command surface and a control surface, whose pivotings about the axisof articulation are commanded, by at least one driven actuator, theconvertible aircraft comprising a self-driven active anti-vibrationsystem with at least one active vibration control computer commandingsaid at least one driven actuator causing said orientable surface topivot about said axis of articulation at least at a frequency of theorder of KbΩ, where b and Ω are, respectively, the number of blades andthe frequency of rotation of each rotor, and K is a whole number atleast equal to 1, so as to at least attenuate, at the fuselage, thefixed wings and the empennage and stabilizer, at least the vibrationsgenerated naturally by the rotation of each rotor.
 3. The convertibleaircraft as claimed in claim 2, wherein said at least one drivenactuator transiently commands the pivoting of said orientable surface ata frequency below KbΩ and of the order of 4 to 6 Hz so as to counter thewhirl flutter phenomenon.
 4. The convertible aircraft as claimed inclaim 2, wherein said at least one driven actuator transiently commandsthe pivoting of said orientable surface at a frequency below KbΩ and,generally of the order of 4 Hz, so as to counter the tail shakephenomenon.
 5. The convertible aircraft as claimed in claim 2, whereinsaid actuator is an excitation ram shaved in movement, maneuvering saidorientable surface against the action of static and dynamic tuningelastic means and driven automatically by said at least one activevibration control computer which is self adaptive and drives said ram onthe basis of signals received from at least one sensors, including load,accelerometer and gyrometer sensors, arranged at least at onepredetermined points on at least one of said fuselage, rotors empennageand stabilizer.
 6. The convertible aircraft as claimed in claim 5,wherein the elastic means absorb the static forces of said orientablesurface and, in dynamic terms, their stiffness is coupled to the inertiaof the moving assembly comprising at least said orientable surface andmoving parts of said ram so as to create a second-order resonant system,the resonant frequency of the moving assembly being tuned to theexcitation frequency of the ram.
 7. The convertible aircraft as claimedin claim 6, wherein said moving assembly has a resonant frequency${f = {\frac{1}{2\quad \pi}\sqrt{\frac{k}{I}}}},$

where k is the stiffness of the elastic means and I is the inertia ofthe moving assembly and the excitation frequency of the ram is nonormally tuned to bΩ such that bΩ is substantially equal to f.
 8. Theconvertible aircraft as claimed in claim 7, wherein the two rotors arethree-bladed rotors, and the excitation frequency of the ram is normallytuned substantially to a frequency of about 20 Hz.
 9. The convertibleaircraft as claimed in claim 5, wherein in that the excitation frequencyof the ram is transiently tuned to a frequency of about 4 Hz to about 6Hz when said sensors detect signals that bear witness to at least one ofthe phenomena of tail shake and whirl flutter and then tuned backsubstantially to the frequency bΩ.
 10. The convertible aircraft asclaimed in claim 5, wherein the excitation ram can also be driven bypilot controls.
 11. The convertible aircraft as claimed in claim 10,wherein command of the excitation ram by the vibration control computeris neutralized while said excitation ram is being commanded by the pilotcontrols.
 12. The convertible aircraft as claimed in claim 2, whereinsaid outboard wing portion is an outboard wing part that can entirelypivot about the axis of articulation and constitutes an elevon.
 13. Theconvertible aircraft as claimed in claim 2, wherein said pivoting partof the outboard wing portion is a pivoting trailing edge elevon of afixed and outboard wing portion.
 14. The convertible aircraft as claimedin claim 2, wherein each fixed wing comprises at least one inboardsurface which is at least one of a command surface and a control surfacebetween the corresponding nacelle and the fuselage and which pivotsabout a second axis of articulation substantially transversal to theaircraft.
 15. The convertible aircraft as claimed in claim 2, whereinthe fixed wings are high wings secured to the upper part of thefuselage.
 16. The convertible aircraft as claimed in claim 15, whereinthe fixed high wings have an upward dihedral angle between the fuselageand the nacelles.
 17. The convertible aircraft as claimed in claim 16,wherein outboard of the nacelles, said outboard wing structure portionsalso have an upward dihedral angle.
 18. The convertible aircraft asclaimed in claim 16, wherein outboard of the nacelles, said outboardwing structure portions have a zero dihedral angle.
 19. The convertibleaircraft as claimed in claim 16, wherein outboard of the nacelles, saidoutboard wing structure portions have a downward dihedral angle so thatthe fixed wing structure of the aircraft is substantially in the form ofgull wings.
 20. The convertible aircraft as claimed in claim 2, whereinthe axis of articulation of the pivoting parts, of the outboard wingstructure portions (13, 13′, 27) is offset forward or blackward from thecenters of twist of the fixed wings.